Conduction pedestals for a gas turbine engine airfoil

ABSTRACT

An airfoil for a gas turbine engine includes an airfoil which defines a leading edge cavity and a forward cavity between a pressure side wall and a suction side wall, the leading edge cavity at least partially defined by a leading edge wall which extends between the pressure side wall and the suction side wall. A rib between the pressure side wall and the suction side wall separates the forward cavity and the leading edge cavity. A pedestal extends between the leading edge wall and the rib.

BACKGROUND

The present disclosure relates to a gas turbine engine, and moreparticularly to an airfoil cooling arrangement.

A gas turbine engine includes a compressor section that compresses airthen channels the compressed air to a combustor section wherein thecompressed airflow is mixed with fuel and ignited to generate hightemperature combustion gases. The combustion core gases flow downstreamthrough a turbine section which extracts energy therefrom to power thecompressor section and a fan section. Since the combustion core gasesare at a high temperature, turbine vanes and turbine blades within theturbine section may have relatively high heat loads at the leadingedges.

SUMMARY

An airfoil for a gas turbine engine according to an exemplary aspect ofthe present disclosure includes a pressure side wall and a suction sidewall which define a leading edge cavity and a forward cavity between thepressure side wall and the suction side wall, with the leading edgecavity at least partially defined by a leading edge wall which extendsbetween the pressure side wall and the suction side wall. A rib betweenthe pressure side wall and the suction side wall separates the forwardcavity and the leading edge cavity. A pedestal extends between theleading edge wall and the rib.

An airfoil for a gas turbine engine according to an exemplary aspect ofthe present disclosure includes a multiple of pedestals which extendbetween a leading edge and a rib, the multiple of pedestals arrayedalong a length of the airfoil between a first end portion and a secondend portion.

BRIEF DESCRIPTION OF THE DRAWINGS

Various features will become apparent to those skilled in the art fromthe following detailed description of the disclosed non-limitingembodiments. The drawings that accompany the detailed description can bebriefly described as follows:

FIG. 1 is a general schematic partial fragmentary view of an exemplarygas turbine engine embodiment for use with the present invention;

FIG. 2 is a perspective view of a vane;

FIG. 3 is a sectional view of an airfoil;

FIG. 4 is a perspective partial fragmentary view of an airfoil with animpingement flow leading edge;

FIG. 5 is a perspective partial fragmentary view of an airfoil with aradial flow leading edge;

FIG. 6 is a sectional view of a leading edge of an airfoil with apedestal according to one non-limiting embodiment;

FIG. 7 is a sectional view of a RELATED ART airfoil leading edge whichillustrates a temperature gradient therein to determine an associatedconduction path axis;

FIG. 8 is a sectional view of a RELATED ART airfoil leading edge whichillustrates a temperature gradient therein to locate the pedestals ofFIG. 7;

FIG. 9 is a sectional view of the airfoil leading edge of FIG. 6 whichillustrates a temperature gradient therein as reduced due to thepedestals;

FIG. 10 is a sectional view of a leading edge of an airfoil withpedestals according to one non-limiting embodiment; and

FIG. 11 is a sectional view of a RELATED ART airfoil leading edge whichillustrates a temperature gradient therein to determine associatedconduction path axes to locate the pedestals of FIG. 10.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 10 which generallyincludes a fan section 12, a compressor section 14, a combustor section16, and a turbine section 18. Within and aft of the combustor section16, engine components are typically cooled due to intense temperature ofthe combustion core gases. While a two spool high bypass turbofan engineis schematically illustrated in the disclosed non-limiting embodiment,it should be understood that the disclosure is applicable to other gasturbine engine configurations.

At least some stages of the turbine rotor blades 22 and turbine statorvanes 24 within the turbine section 18, for example, may be cooled witha cooling airflow typically sourced with a bleed airflow from thecompressor section 14 at temperature lower than the core gas within theturbine section 18. The cooling airflow passes through at least onecooling circuit flow path 26 (FIG. 2) to transfer thermal energy fromthe component to the cooling airflow.

Each cooling circuit flow path 26 may be disposed in any component thatrequires cooling, and in most cases the component receives coolingairflow therethrough as the external surface thereof is exposed tocombustion core gases. In the illustrated embodiment and for purposes ofgiving a detailed example, the cooling circuit flow path 26 will bedescribed herein as being disposed within a portion of an airfoil 32such as that of a stator vane 24 or rotor blade 22. It should beunderstood, however, that the cooling circuit flow path 26 is notlimited to these applications and may be utilized within other areassuch as liners, seals, and other structures with stagnation regionsexposed to high temperature core gas flow.

With reference to FIG. 2, the cooling circuit flow path 26 communicateswith a multiple of cavities, for example 34A-34B shown in FIG. 3, formedwithin the airfoil 32. The multiple of cavities 34A-34B direct coolingairflow which may include air received from the compressor section intohigh temperature areas of the airfoil 32.

The airfoil 32 is defined by an outer airfoil wall surface 40 between aleading edge 36 and a trailing edge 42. The outer airfoil wall surface40 typically has a generally concave shaped portion forming a pressureside 40P and a generally convex shaped portion forming a suction side40S which are connected by a leading edge wall 40L at the leading edge36. The outer airfoil wall surface 40 is longitudinally defined to spana first end portion 46 and a second end portion 48. The end portions 46,48 may include features to mount the airfoil to other structures such asengine static structure or rotor disk. For example, the end portions 46,48 for a vane may include outer vane platforms and for a blade mayinclude an attachment section and a blade tip. It should be understoodthat various component arrangement may likewise be utilized with thepresent invention.

With reference to FIG. 3, the forward cavity 34A is generally defined bya first rib 54 just aft of the leading edge 36. The first rib 54separates the forward cavity 34A from a leading edge cavity 56 definedat least partially by the outer airfoil wall surface 40 and oftenreferred to as a “peanut” cavity. The first rib 54 may, for example, atleast partially define an impingement leading edge 62 (FIG. 4) or aradial flow leading edge 64 (FIG. 5) which may span a portion of or theentire length of the airfoil 32. That is, the pedestals 60 may bespecifically located along the entire airfoil 32 span or a selectportion or portions thereof.

The leading edge cavity 56 includes the multiple of pedestals 60 whichare transverse to and extend between the leading edge 36 and the firstrib 54. It should be understood that any number of pedestals 60 may beso positioned. The pedestals 60 provide an additional thermal conductivepath along a conduction path axis H (FIG. 6) from the leading edge 36 tothe first rib 54 to reduce the temperature of the leading edge 36 as theleading edge 36 may otherwise be hundreds of degrees hotter than thepressure side 40P and suction side 40S of the airfoil 32 due to higherexternal heat transfer coefficients at the stagnation region S (FIG. 7).It should be understood that the stagnation region S is a region withinwhich the combustion gas flow Mach number may be relatively low suchthat a temperature concentration occurs.

For the impingement leading edge 62 cooling scheme (FIG. 4) the firstrib 54 may define a multiple of cooling holes 66 which communicate acooling flow from the forward cavity 34A into the leading edge cavity 56through the first rib 54 then out through a multiple of leading edgecooling holes 68. That is, the cooling flow is communicated generallyalong the pedestals 60. For the radial flow leading edge 64 coolingscheme (FIG. 5) the cooling flow from within the leading edge cavity 56passes transverse to the pedestals 60 and out through a multiple ofleading edge cooling holes 70. It should be understood that various suchcooling schemes will benefit from the pedestals 60.

The pedestals 60 reduce leading edge 36 temperatures mainly from theenhanced conduction effects of the pedestals 60 from the leading edge 36to the first rib 54 (FIGS. 8 and 9). In addition, for radial flowleading edges (FIG. 5), a portion of the metal temperature reduction isachieved by the enhancement of the internal heat transfer coefficient ascoolant flow passes over the pedestals 60. The lower temperature at thestagnation region beneficially results in, for example, a higheroxidation, local creep, and Thermal Mechanical Fatigue (TMF) capability.

The pedestals 60 may be selectively oriented at a multiple of differentangles in the leading edge cavity 56 to achieve the desired thermalreduction effect. That is, the pedestals 60-1, 60-2 may be aligned alongconduction path axes H1, H2 (FIG. 10) which extend into the highesttemperature areas in the stagnation region of the leading edge 36 (FIG.11) to facilitate a more direct heat transfer from the leading edge 36to the first rib 54. It should be understood that the axes H1, H2 maychange along the span of the airfoil 32. The relative positions of thepedestals 60-1, 60-2 may thereby also change along the span tocorrespond therewith.

The manufacture of the pedestals 60 may be achieved by a proprietaryFugitive Core Process which uses thermoplastic inserts to create a onepiece core with multiple pull angles as developed by Alcoa Howmet ofCleveland Ohio USA. Generally, sacrificial thermoplastic pieces make upthe rib and leading edge pedestals; the thermoplastic pieces areassembled into the core die and core material is injected around thethermoplastic pieces; the thermoplastic pieces are melted, leaving voidsin finished core; and metal fill voids in core to form pedestals in thefinished part.

It should be understood that like reference numerals identifycorresponding or similar elements throughout the several drawings. Itshould also be understood that although a particular componentarrangement is disclosed in the illustrated embodiment, otherarrangements will benefit herefrom.

Although particular step sequences are shown, described, and claimed, itshould be understood that steps may be performed in any order, separatedor combined unless otherwise indicated and will still benefit from thepresent disclosure.

The foregoing description is exemplary rather than defined by thelimitations within. Various non-limiting embodiments are disclosedherein, however, one of ordinary skill in the art would recognize thatvarious modifications and variations in light of the above teachingswill fall within the scope of the appended claims. It is therefore to beunderstood that within the scope of the appended claims, the disclosuremay be practiced other than as specifically described. For that reasonthe appended claims should be studied to determine true scope andcontent.

1. An airfoil for a gas turbine engine comprising: a pressure side walland a suction side wall which define a leading edge cavity and a forwardcavity between said pressure side wall and said suction side wall, saidleading edge cavity at least partially defined by a leading edge wallwhich extends between said pressure side wall and said suction sidewall; a rib between said pressure side wall and said suction side wallto at least partially divide said forward cavity and said leading edgecavity; and a pedestal which extends between said leading edge wall andsaid rib.
 2. The airfoil as recited in claim 1, wherein said pedestal isaligned along an axis which extends toward a high temperature area in astagnation region of said leading edge wall.
 3. The airfoil as recitedin claim 2, further comprising a second pedestal aligned along a secondaxis different than said axis.
 4. The airfoil as recited in claim 1,wherein said rib at least partially defines an impingement leading edge.5. The airfoil as recited in claim 4, wherein said rib defines amultiple of cooling holes which communicate a cooling flow from saidforward cavity into said leading edge cavity through said rib thenthrough a multiple of leading edge cooling holes through said leadingedge.
 6. The airfoil as recited in claim 1, wherein said rib at leastpartially defines a radial flow leading edge.
 7. The airfoil as recitedin claim 6, wherein said leading edge defines a multiple of coolingholes which communicate a cooling flow from within said leading edgecavity through a multiple of leading edge cooling holes through saidleading edge.
 8. The airfoil as recited in claim 1, wherein said airfoilat least partially defines a turbine vane.
 9. The airfoil as recited inclaim 1, wherein said airfoil at least partially defines a turbineblade.
 10. An airfoil for a gas turbine engine comprising: a pressureside wall and a suction side wall which defines a leading edge cavityand a forward cavity between said pressure side wall and said suctionside wall, said leading edge cavity at least partially defined by aleading edge wall which extends between said pressure side wall and saidsuction side wall; a rib between said pressure side wall and saidsuction side wall to at least partially divide said forward cavity andsaid leading edge cavity; and a multiple of pedestals which extendbetween said leading edge wall and said rib, said multiple of pedestalsarrayed along a length of said airfoil between a first end portion and asecond end portion.
 11. The airfoil as recited in claim 10, wherein eachof said multiple of pedestals are aligned along an axis which extendstoward a high temperature area in a stagnation region of said leadingedge wall.
 12. The airfoil as recited in claim 10, wherein a first setof said multiple of pedestals are aligned along a first axis whichextends toward a first high temperature area in a stagnation region ofsaid leading edge and a second set of said multiple of pedestals arealigned along a second axis which extends toward a second hightemperature area in the stagnation region of said leading edge.
 13. Theairfoil as recited in claim 10, wherein each of said multiple ofpedestals are transverse to said rib.
 14. The airfoil as recited inclaim 10, wherein said airfoil at least partially defines a turbinevane.
 15. The airfoil as recited in claim 10, wherein said airfoil atleast partially defines a turbine blade.